A gas turbine generally includes a compressor, a combustor disposed downstream form the compressor and a turbine section disposed downstream from the combustor. A working fluid such as air enters the compressor where it is progressively compressed to provide a compressed working fluid to the combustor. Fuel is mixed with the compressed working fluid within the combustor and the mixture it is burned to produce combustion gases at a high temperature and a high velocity. The combustion gases are then routed from the combustor into the turbine section where thermal and/or kinetic energy are extracted to produce work.
The turbine section generally includes a plurality of rotor blades that extend radially from a rotor disk that is coupled to a rotor shaft. The rotor blades are circumferentially surrounded by a casing. Each rotor blade includes a blade tip that is defined at a distal or radial end of the rotor blade. A shroud assembly extends circumferentially within the casing around the plurality of rotor blades. The shroud assembly is typically mounted to an inner surface of the casing. The shroud assembly often comprises a number of shroud block segments that are arranged in an annular array around the tips of the rotor blades.
The plurality of rotor blades and the shroud block segments at least partially define a hot gas path for routing the hot combustion gases through the turbine section. A small radial gap is generally defined between the blade tips and a hot side portion of the shroud block segments. The radial gap is designed or sized to provide radial clearance between the blade tips and the hot side portion of the shroud block segments, while also providing a partial fluidic seal to control leakage of the combustion gases over the blade tips during operation. Leakage of the combustion gases over the blade tips generally results in a decrease in overall turbine efficiency.
The rotor blades and shroud block segments, particularly the hot side portions, are subjected to the high temperature combustion gases as they flow through the turbine section. As a result, cooling of the rotor blade tips and the shroud block segments is necessary to reduce thermal stresses and to improve durability of those components. One cooling scheme for cooling shroud block segments includes directing a cooling medium such as a portion of the compressed working fluid onto a backside portion of each shroud block segment. The cooling medium is routed from the back side portion into a cooling channel that is defined within the shroud block segment via a plurality of cooling passages. The cooling medium is then exhausted into the hot gas path via one or more exhaust passages defined the shroud block segments. The cooling channel is in thermal communication with the hot side portion, thereby allowing for heat transfer between the hot side portion and the cooling medium before the cooling medium is exhausted from the cooling channel.
The cooing passages are generally machined and/or cast into the shroud block segments. Once the cooling passages have been cast and/or machined into the shroud block segment the ability to later modify the size, pattern and quantity of the cooling passages thereby modifying or tuning the cooling provided to the shroud block segment becomes limited. Therefore, a system for cooling a shroud block segment which provides for cooling flow flexibility would be useful.